text:
How dissimilar are redundant flight control computers?
Facts
On Airbus aircraft there are computers to secure the flight envelope, or to move the control surfaces. FADECs totally control the engines. Computers take decisions in place of the pilots, or even against their commands. Boeing aircraft have similar computers, even if the crew has more authority.

A330 electronic bay, source, photo by 'swiss_a320'
Safety implications
Being critical, systems are redundant and supervise each other to detect possible failures and isolate failed components. However a computer can still be developed from faulty specifications, or be wrongly manufactured, and a program can contains bugs. If the same defect is present on all computers on the production line, the purpose of redundancy may be defeated, as they wouldn't be able to detect erroneous behavior.
This is better stated in this article:

Because of the severe consequences resulting from a single point of
  failure, hardware redundancy is critical in DAL A systems. But if the
  aircraft uses a redundant architecture built with similar channels,
  that system will still be susceptible to common mode failures that can
  cause all channels to fail in the same way.

Question
What are the principles used in aviation to reduce the possibility for redundant computers to fail or to make the same errors at the same time?

3 or 4, usually. But to answer your other question: if all redundant computers are manufactured or programmed in the same faulty way then the crew and passengers would indeed be screwed if it made through all reviews, auditing, and testing and be found on a operational product, which is why making such a thing and get it through the process and getting certificated is very expensive and time consuming.

But, unfortunately your concern is legit and a lot of test pilots have lost their lives because of faulty computer and computer programs, and I guess so have passengers and crew members, too.

Very much related (do read beyond the title): Why are critical flight computers redundant?

Example of a bug making it through testing. Fortunately it’s quite a rare thing.

How much weight are you willing to sacrifice?  What penalty are you willing to pay for the next level of redundancy?  Have you briefed the CEO?  Do you want to be employed tomorrow morning?

Perhaps they make the programmers go up on the first test flight.

@Notts90: related issue regarding Patriot missile software bug in the 1991 Gulf War.

There's two separate issues here, redundancy and dissimilarity.  Level of redudancy is just a measure of how many systems you have commanding and monitoring, while dissimilarity is how separate and independent those system are.  I'd suggest clarifying which one you're curious about by renaming the title something like "How dissimilar are redundant flight control computers"

@CodyP: Thanks, that's a good suggestion.

@mins I'm a little divided on the title change. It feels slightly like the question has been changed, and invalidates parts of some answers (parts 2 and 4 of the accepted answer, for example, aren't about computers). Maybe the old title and a mention within the question of "how are redundant systems made dissimilar"? Or maybe just removing "computers", because a lot of the redundancy is outside the computers themselves. I'm not sure... I'll leave it to you to decide.

The duplicated systems are built on different chips, using different programming languages (ADA is one), and coded by different teams to the same functional specification. This minimizes the chance of the same operating parameters (e.g. airspeed, angle of attack etc.) producing the same bug-driven 'error' in two systems. After AF447, Airbus made changes to the software systems to simplify switching one off where there are conflicting results.

@Pete855217 - you seem to have specifically and actually answered this great and fascinating question.  Thanks for that

BTW the image in the question is ...... somewhat terrifying, in a way!

As far as Airbus is concerned:

Each unit is composed of two dissimilar boards, one driving the output and the other checking it. Dissimilar means both different CPUs and chipsets (A320 uses i386 (Intel) and m68k (Motorola); newer models use different combinations, basically whatever was widely used at the time they were designed) and software written by two independent teams.
There are fail-overs, two or three depending on the system (IIRC the unit reading the side-sticks is the only one with four copies).
The two main axes, pitch and roll, are controlled by two different systems. ELAC controls elevator and ailerons, SEC controls horizontal stabilizer and spoilers. This is two completely independent chains including different actual control surfaces except for the side-sticks.
A320 has (hydro)mechanical backup for pitch via the trim wheel and yaw via the pedals, utilizing yaw-roll coupling for roll. This works even with complete electrical failure. IIRC the backups on newer models don't though (because complete electrical failure has never happened).

Expanding on the dissimilar boards bit - idk about airbus or anyone else in particular, but I have heard of having three different systems for a decision - this means that if one is "faulty", it is not in agreement with the others, and the correct decision can be upheld, but you can't tell which is which with two systems.  Is this right?

@Baldrickk Yes, that's the normal reason for two-out-of-three voting: with two, if one is wrong, you can't know which one is wrong; all you can know is that they disagree. With three, if one is wrong, you can know because two are still in agreement. As long as not two out of three are wrong, three gives you far better ability to handle disagreements than do two.

"because complete electrical failure has never happened" >> You'd think aircraft manufactures would have learned about Murphy's Law long ago. Ask Al Haynes about "impossible" failures. https://en.wikipedia.org/wiki/United_Airlines_Flight_232  :-/

@MichaelKjörling See: Minority Report

@Baldrickk, In general yes, but it never applies to Airbus. All airbus systems have one unit making the decision and another one system verifying it. And if the verification fails, the two-system unit is declared faulty and a fail-over approach is used.

"...because complete electrical failure has never happened" - gosh, I hope I'm not on the first plane this happens to.

@Shawn, that's the thing. Hydraulic failure did happen, but it is needed for the flight control anyway. Since electric failure is less likely than hydraulic one, requiring electric power does not increase the risk that much.

Understood, and I'm not disagreeing with you. However, it is very short-sighted of Airbus to think that an extremely improbable situation is impossible. Again, ask Capt. Haynes about what's impossible. All sorts of "impossible" things happened that day.

@JanHudec thanks for the additional info.  it's interesting to know that they would do it that way.

@Shawn: you seem to be implying that you know better than Airbus regarding flight safety. Could you expand on that?

@MartinArgerami I apologize if that's how you took what I wrote. I'm simply saying that there are some pretty good real-world examples of things happening that the aircraft manufacturers thought were impossible - like a total hydraulic failure on a DC-10. It probably wouldn't hurt Airbus to keep those examples in mind if they want to deal in "impossible" situations. And in my experience, flying (and other things) always included being prepared for the impossible and improbable so that the likely stuff, no matter how hard, was much easier to handle.

@Shawn: no need to apologize. All I wanted to imply is that there is a lot of knowledge on fly safety. Both Airbus and Boeing have been building planes for a long time, and the knowledge on accidents has increased year after year. Reality is that we are already at the stage where most accidents are not caused by design failures.

Redundancy is not only achieved by multiplying the computers, but also by diversifying them. On Airbus airliners, two different computers are used (one with Intel chips, the other with Motorola chips in case of the A320) and software is written twice, one for control, the other for monitoring, by two teams which are not allowed to interact.
To cite from chapter 12 of The Avionics Handbook:

Despite the nonrecurring costs induced by dissimilarity, it is
  fundamental that the five computers all be of different natures to
  avoid common mode failures. These failures could lead to the total
  loss of the electrical flight control system. Consequently, two types
  of computers may be distinguished:
2 ELAC (elevator and aileron
  computers) and 3 SEC (spoiler and elevator computers) on A320/A321
  and,
3 FCPC (flight control primary computers) and 2 FCSC (flight control
  secondary computers) on A330/A340.
Taking the 320 as an example, the
  ELACs are produced by Thomson-CSF around 68010 microprocessors and the
  SECs are produced in cooperation by SFENA/Aerospatiale with a hardware
  based on the 80186 microprocessor. We therefore have two different
  design and manufacturing teams with different microprocessors (and
  associated circuits), different computer architectures, and different
  functional specifications. At the software level, the architecture
  of the system leads to the use of four software packages (ELAC control
  channel, ELAC monitor channel, SEC control channel, and SEC monitor
  channel) when, functionally, one would suffice.

Famously the 777 has three redundant computers created by different teams/companies in different programming languages.

In my experience sometimes even the underlying algorithms have to be different (if possible).

Holy carp! The Motorola 68010 and Intel 80186? Those are positively ancient!! Though, I guess when the A320 first rolled off the line in 1987 they were pretty new...

@FreeMan: The Eurofighter uses also the Motorola 68000. Barring some midlife upgrade, those have to be available for the next 30 years ore more. To keep the last one flying will probably require to steal the chips from a museum. At least the more modern Airbus computers use the PPC.

That begs a whole new question (which may not be appropriate for this forum) of how does Airbus keep Intel & Moto producing these ancient CPUs and/or how many do they (their supplier) have stockpiled? Intel & Moto have long since recouped their dev & startup costs, so producing more should be nearly pure profit, but it costs a lot to keep fab plants running for tiny quantities,. /OT musings

@FreeMan: Those are military grade chips which have passed a very long approval process. Part of the process is to make sure that somebody will be able to manufacture those chips in decades to come. Of course with plenty of government financial support. These are really juicy contracts, so the manufacturers will not die out or lose interest. Hint: It's not alone the toilet seats that are more expensive than their equivalent in massive gold.

@FreeMan Even if the cost of a mil-grade 68000 was a "ridiculously high" number like $10,000 per part, that would be negligible compared with the cost of certifying a newer chip design. For example even one hour of flight test time would cost more than $10,000 just to fly the aircraft, not counting the cost of actually monitoring the behaviour of anything inside it.

... some mandatory engine certification procedures have budgets of the order of 20 or 30 million USD - and double or triple that if the tests are not passed first time and need to be re-done. That sort of money would pay for a lot of CPU chips! But even USD 20m is a fairly small number in a complete "new engine" project with a budget of the order of USD 1bn.

In general, software isn't manufactured wrong. When the software is created (programmed), defects can be introduced as you described by either faulty implementations or by bad specifications. Faulty implementations are detected by testing the software. Testing takes many forms; unit testing is one of the more basic forms, where individual functions of the underlying programming code is tested to see if it is implemented correctly. This can scale upwards when doing system and integration testing where larger pieces of the software is coupled together to see how it performs as a whole. But simply testing the code at this level doesn't catch everything. Writing a program is rarely about getting it to do what you want it to do, it's mostly about handling all the strange edge-cases and failure scenarios. And this is where most software fail.
To guard against such cases, you can run through audits, simulations, static code analysis and lots of other forms of inspections and testing.
Faulty specifications is a different beast, where you have to rely on documentation. In a perfect world each requirement must be documented to a level describing why the requirement exists, and any input and output that should result from it if applicable. Specifications are developed by multiple people to guard against one person forgetting something, or wrongly interpreting something, but this doesn't catch everything either.
To add another level of protection against software defects, you add multiple instances of the system, and you also have a team create their own version of the systems, preferably on different hardware. You can then divide responsibility of certain subsystems and spread it out among the various computers running the system, adding another level of redundancy as well as to lessen the computational load on each computer, and the risk that any parts of the system interact in unforeseen ways.
The Fast Company had an excellent writeup on the process of writing software for the space shuttle. Although it isn't directly related to neither Airbus or Boeing, it gives an insight into how the process worked and what it resulted in.

The software for the Space Shuttle was also one of the most expensive software projects ever, in terms of almost every measurable metric. There can be no denying, however, that it did produce quality that most software projects can never hope to even approach (and I say this as a software developer myself).

I knew a guy who worked on the shuttle software. I don't think he ever quite recovered from it...

@BobJarvis I'm curious - could you share details?

There is a fundamental difference between Space Shuttle and Airbus. Space Shuttle had one implementation of the software and one implementation of the hardware, relying on careful implementation and only using fail-over for in-operation mechanical failures. On the other hand Airbus uses multiple implementations and relies on fail-over for handling both mechanical failures and software and hardware design bugs.

@JanHudec: Actually, the space shuttle had one implementation of the hardware, but two implementations of the software (the primary and backup flight software, developed by two separate teams based on the same set of specifications).

Primary systems should have identical computers and software and is the case on many airborne vehicle systems computers. However, the independent backup systems should have dissimilar systems and software depending on the architecture and redundancy management requirements and schemes for safety. Other than flight controls which do have dissimilarity in hardware, primary flights displays for pilot and copilot sides for airspeed and inertial navigation are often triple redundant to retain the attitude function. These correctly use identical 3 nav system computers whereas the "backup" is dissimilar for purpose of flight safety critical functions and determinism. The overall system architecture of parallel or more (triplex) must have independent and redundant systems that meet the agency and regulatory criteria for safety and airworthiness as well as reliability and availability. Generally, having identical computers for "primary systems will require in depth fault insertion testing of combination of complex interaction will minimize the possibility of software faults and sometimes unfounded fears that defects will somehow manifest in latency. Proper testing in all environments is the key to getting rid of any defects that would cause potential hazards and risk. Software safety methods are recommended to prevent, eliminate and control such issues to ensure safety and airworthiness requirements are met. Safety analyses and independent reviews are required in these cases with approvals.

authors: Bob Jarvis - Слава Україні


text:
Why is the horizontal stabilizer root interface designed this way?
Most large aircraft have a similar design for their trimmable horizontal stabilizer: they connect to the tail cone on a flattened area which allows them to be rotated within some range:

From left to right: A319 (source), A320 (source, B787 (source)
But the THS of the A330 is different, it looks like a frog eye... What is the purpose of this flexible wall?

Source

It is a seal- to prevent entry of air into the stabilizer mounting area. For aircraft with trimmable horizontal stabilizers (THS), there is an opening to the rear of the aft bulkhead to allow for the movement of the THS. It can be seen in the following image for Boeing 787.

Boeing 787 aft fuselage structure; image from aero-news.net
Now, as the opening will be there while the THS moves, a seal has to be present to prevent entry of air (and FoD etc.). Boeing went ahead and put a hinge aperture sealing plate over that area, with only a small opening where the actuated area of the THS moves, with the seal inside the structure, as can be seen below.

777 THS area; image from deicinginnovations.com
Without cover, the are will look something like this:

Trimmable horizontal stabilizer region without cover, ; image from skysoftairlines.blogspot.in
Airbus, on the other hand went ahead and put a tailplane sealing plate outside the opening. This was used in the A330 and initial versions of A340 (which shared the fuselage with A330), before Airbus changed its mind and made the THS region similar to the other aircraft in A340-500 and -600 versions, where the horizontal stabilizer was redesigned.
Also, see these (1 & 2) patents and this discussion.

This structure prevents air from flowing between stabiliser and fuselage, which helps to increase tail effectivity. In this particular case the A330 has a wider trim angle range which produces more vertical travel of the root of the stabiliser which in turn requires a more elaborate seal.
As any other part of the outer shape of an aircraft, the intersection between fuselage and horizontal tail is optimised for performance and construction cost. Where the compromise between both is struck depends on the trim angle range, the manufacturer's philosophy and its culture.
If you look at the tail from above, you will see that the fuselage is coke-bottled there in order to smooth out the cross section distribution over length, like it is done on supersonic fighters. This produces a flat surface on the side just where the biggest thickness of the horizontal tail sits. Unfortunately, this is not where the largest vertical travel of the root of the tail surface happens; this is either at the leading edge (for highly swept tails) or at the trailing edge.
For the ERJ-170, this flat surface is sufficient to cover the travel of the stabiliser nose, but for the A330 a wider range of trim angles is required. This might be because Airbus wanted a wider cg range or because the flaps produce a higher pitching moment change.

source: Stack Exchange

license: Creative Commons - Attribution Share-Alike - https://creativecommons.org/licenses/by-sa/4.0/

all_licenses: Creative Commons - Attribution Share-Alike - https://creativecommons.org/licenses/by-sa/3.0/

sort: votes


text:
Estimating Moment Coefficient about Aerodynamic Center for wing-body combination
Are there any references or analytical methods to estimate the moment coefficient about the aerodynamic center, C_M,ac for a given wing-body combination? Even something rough would be useful such as "adding fuselage to the wing increases magnitude of CMac by ~x%". If anything, I'm looking for a conservative estimate that overpredicts C_M,ac for some 1st principals structural calcs.
I understand that C_M,ac for the wing alone can be obtained from XFOIL/airfoil databases. Adding a fuselage will contribute a negative increment to C_M,ac. The resulting C_M,ac for wing-body is usually measured in a wind tunnel.
Wind tunnel measurements are used to characterize interference effects between the wing and fuselage, which are difficult to predict theoretically. Because of this, the moment due to the wing-body combination isn't simply just the sum of the separate wing and fuselage moments. This being understood, I wonder if the sum of individual moments would give a conservative estimate for C_M,ac that I can use.

In this case the NACA report 711 is your friend.
It gives a very easy empirical equation to evaluate the fuselage's moment coefficient based on flight tests. The report is quite old but if your fuselage has a conventional shape then the value given by this equation should be reasonable.
$\frac{dC_m}{d\alpha}=\frac{k_fw²_fL_f}{cS}$
where:

$\frac{dC_m}{d\alpha}$ is the moment coefficient's derivative, i.e. the variation of the moment coefficient in respect to $\alpha$;
$k_f$ is an empirical value given by the following plot;
$w_f$ is the fuselage's width;
$L_f$ is the fuselage's length;
$S$ is obviously the wing surface;
and $c$ the mean aerodynamic chord.

Units of measurement are feet and degrees.

Thank you sophit - this is very helpful. I also found NACA Report No. 540 to be super useful as well (one of the references in Report 711)

@Seth: I'm glad I could help :)

all_licenses: Creative Commons - Attribution Share-Alike - https://creativecommons.org/licenses/by-sa/4.0/


text:
Fuselage Bending Stress Calculation Approach
Does the approach below seem reasonable for a first-order estimate of the bending stress on a fuselage? I want to specifically confirm the end result, which shows that the maximum bending moment, and therefore maximum stress occurs at the nose.
Approach/Assumptions:

Model fuselage as a beam in bending
Beam is fixed at the CG, with lift from the wing and tail acting perpendicular to the beam
Weights of each component - battery, wing, payload, motors are combined into a singe load that acts thru the CG of the airplane
Thrust and drag forces act horizontally thru the fuselage centerline

Calculations:

Update
After following sophit's suggestion to split up the weight of each component and fix the beam at the wing, I'm now getting that the max bending moment occurs near the wing, which makes more sense.

I'd fix the fuselage at the wing (point A) instead at the CG; and I'd put the weight of each component where each component is instead of concentrating everything in the CG. Max bending moment at the nose looks strange to me... maybe max deflection?

One problem with you current iteration of calculation is, that the lift of the wing is ahead of the CG. This causes the problem with a moment existing at the nose of your aircraft since the whole systems moment isnt balanced.

Thanks for the input sophit. I ended up splitting up the weights of each component and fixing the fuselage at the wing. I'm now getting a more reasonable result with the max bending moment at the wing. See updated shear and bending plots in the post.

Thanks for the input hph304j. That is a good point - I'm estimating the CG to be fwd of the wing, this was just an error in the diagram. Since I ended up splitting the weights of each component, the precise CG location was not used for the calcs

I'm glad I could help :-)  Now it definitely looks more realistic


text:
Wing Structural Load Test Simplification
I'd like to perform a static structural load test of a wing for a UAV. The wing is a rectangular planform of constant thickness throughout the wingspan with no twist, dihedral, or sweep.
During steady-level flight, the lift distribution will be a uniformly distributed load. In order to simplify the test, I'd like to load the wing with a concentrated force at the tip as opposed to a distributed load along the span. This allows me to use a single weight as opposed to various sandbags.
Are there any issues in applying a concentrated load at the wing tip that would result in an equivalent maximum bending moment at the root to verify its structural integrity?
Actual Load::
 where M_max = FL/2
Approximated Load:
 where M_max = FL
Where F_approximated = F_actual/2

Moment and shear are the same in both cases at the root, but obviously their distributions along the wingspan is different.

For sure, but if anything the fixed point load case should be over-conservative compared to distributed load case. So I should be covered if the fixed load test is successful I would think.

In your approximation the moment is a linear function of distance, in the actual load case it is quadratic function. For this reason your approximation is a conservative approximation. I see no problem, but I am no structural engineer

Yes, it is actually conservative... or your wing is oversized and heavier :-)

If your wing is rectangular and has no twist or change in airfoil, then your lift will not be uniformly distributed (I'm assuming you mean the same lift per span), but rather somewhere between rectangular and elliptical, see Schrenk's approximatino method.

I'm with @sophit with this one. If your wing can handle the standard load but loaded entirely at a point at the tip, then the wing panels and spar are significantly stronger-- and heavier-- than needed. Maybe you're fine with that, in which case go for it.

Thanks all for the feedback - agree with your points. For ease of manufacturing, I went with a constant cross-section airfoil along the span, which was sized based on the max bending stress at the root. Understood that it is heavier than ideal, but am trading that for simplified manufacturing.

So to sum up, no one put up an answer but OP ist happy anyways :D

@U_flow: that's the scope of this SE, make people happy :-D


text:
Does the sound barrier apply to a silent aircraft?
For my summer duty, I was looking at the sound barrier effects with a glimmer of hope to understand them:

"A supersonic aircraft is one that can travel faster that the speed of sound. As the plane approaches the speed of sound, it catches up with the sound waves traveling in front of it and pushes them up against each other. The result is a barrier of squashed air in front of the aircraft." Source
Is it theoretically possible to design an aircraft totally silent or at least quiet?
In that case would the problems associated with transonic/supersonic speeds disappear, or at least are reduced?

As the plane approaches the speed of sound, it catches up with the sound waves traveling in front of it and pushes them up against each
  other.

Or is there a silence barrier too?

An airplane makes sound just by running into air.  Much of the noise you associate with a jet plane is not the engine noise, but the noise of the air being disturbed.  A glider at high speed makes a sound not altogether unlike that of a jet.

Related: Would a perfectly silent supersonic aircraft create a sonic boom?

An airplane which disturbs no air at all when it moves will be totally silent.

That graphic seems to reinforce the myth that the sonic boom only occurs at Mach 1.0. It would be nice if you included a graphic showing a 'boom carpet' instead.

The sound barrier has nothing to do with the noise the aircraft makes. It has to do with the fact that if it is going fast enough the air can't get out of the way. This creates a shock wave where the aircraft is pushing the air aside. A shock wave is where the speed of the air transitions from faster than sound to slower than sound and involves the release of considerable energy. You can control the shockwave to some extent by careful shaping of the surfaces, but a plane going fast enough can't be silent.
You can make very quiet sub-sonic aircraft, but they tend to go slowly so that the noise of the air flowing over the craft is minimized.

I wanted to add here that the shuttle produced sonic booms despite having no power during reentry

I'll try to keep this really simple, as OP's question leads me to believe they are not familiar with the physics involved.
As indicated by others here, the sound barrier is not directly related to noise, but rather to the nature of sound itself:

Sound is just a pressure wave moving through the air.
In turn, the speed of sound is the speed at which a disturbance (pressure wave) can propagate through air.
Us hearing these disturbances (or not, there is a lot of sound that escapes us) is just an evolutionary adaptation.
For air to "move aside" to let an object like a plane fly through, it must sense it aproaching. Aircraft disturb air a good distance in front of them, this may sound a bit odd the first time you hear it.
If the plane is faster than the speed of sound, the air cannot "move aside" in the usual manner, as the aircraft outruns the pressure wave that would have moved the air out of the way in time. This is why fast aircraft measure their speed using the Mach number to show how fast they are in relation to sound, as this becomes the most important factor once you approach $M_{\infty}=1.0$
The end effect is that a supersonic aircraft must then force the air out of its way in a more forceful manner, compressing and accelerating it to its own speed. This phenomenon is commonly called the sound barrier.
As an added point of interest, a similar phenomenon happens at the rear of an aircraft flying over the speed of sound, the airflow cannot adapt to the end of the body fast enough, creating a low pressure area.

Air doesn't really "sense" anything. What is happening is a pressure wave is created in front of the aircraft as it pushes the air out of the way. This is called a "bow shock" and is the same principle as a boat pushing water out of the way ahead of the boat.

@Ron Anthropomorphic explanations are usually perfectly acceptable. Air does not sense, but it is disturbed by pressure propogating at the speed of sound - much like you do not sense sound but your eardrums are displaced by the pressure.

As both aircraft displacement and aircraft noise create a pressure wave which propagate at the "speed of sound", I imagine both participate to the creation of the shock wave. What is missing in your answer is in which proportion, and how much the shock wave energy would be reduced if the aircraft sound was reduced. This may be not significant, but that what the question is about.

license: Creative Commons - Attribution Share-Alike - https://creativecommons.org/licenses/by-sa/3.0/


text:
How are electronic systems locked and connected in the electronic bay?
On aircraft the electronics bay contains multiple systems, e.g. for the A320:

A320, photo by B-2230, source: JetPhotos
How is the electronic equipment locked (it seems there are two mobile screws) and how are the wires connected? Details will be appreciated: standards, diagrams, images...

This company page (not affiliated) seems to provide better images of the various trays and fasteners: http://avionicshardwareandmounts.com/

And this company's product catalog seems to provide some illustrations of the hold down mechanisms: https://www.arconic.com/fastening_systems_and_rings/aerospace/catalog/pdf/afs_camloc_ram_holddown.pdf

@selectstriker2: Very good finds, that's more than what I was looking for.

@selectstriker2 I totally think you should add that to your answer, it completes it quite nicely

The LRUs are attached to trays which house the electrical connectors and the cabling behind them:

source
The hold down mechanisms vary in design, but it appears they generally attach the LRU to the tray at the front, the back likely being held in by the electrical connectors at the back.

The LRUs slide into the trays and are electrically mated at the rear. There are small hooks on the bottom front of the LRUs. The hold down is a screw collar on the bottom front of the tray that swings up and is screwed down over the hook. Additionally, the trays support forced-air cooling of the LRUs. Air is forced up through the LRUs (the plastic on the bottom of the trays in the pic is a flow restrictor.) It's all spelled out in ARINC Characteristic 600.


text:
Propeller Curve Modification on Test Rig
this is my first stack exchange question so please bear with me if I have missed a rule: I tried searching for this question to no avail.
I have an off-the-shelf propeller for the purpose of eventual high-altitude flying. I used the airfoil profiles to make an approximate propeller efficiency/c_thrust/c_torque/c_power vs advance ratio curves, as well as the power/thrust/torque/rpm vs airspeed.
My question is: if I had to spin up this prop on a test rig on the ground, there are two major changes I can see - atmospheric air density, and the fact that airspeed is zero. How can I predict how the propeller will behave when the airspeed is held zero (the rig will cancel out the thrust generated). Looking for things like max possible RPM, torque, etc. and how the behavior of the propeller changes.
Thanks a lot, happy to provide more details.

Fortunately, the speed in the propeller plane is not zero once the propeller runs. It will suck in and accelerate the air ahead of it as much as it pushes air out the back, accelerating it further. This was already formulated by Robert Edmund Froude and is called Froude hypothesis. Read this answer for more.
Using the equations from the linked answer you can already calculate your propeller's efficiency and the flow speed in the propeller plane once you run it on your test bench. Most likely your propeller is designed for more than the flow speed achievable with static conditions, so this will help you to calibrate your calculations only for low prop speeds. But since there is a flow speed already, all equations will work.
Your propeller has an advance ratio that tells you how fast it should be spun for a given flight speed. In order to extract more thrust, the propeller should be spun a bit faster, so a positive angle of attack at the blades results over the whole propeller span.
In static conditions, you will not achieve a good angle of attack over the whole span. The faster you spin the propeller, the more it will experience too high an angle of attack, a condition that will be worst at the root and only the tips will exhibit close to proper flow conditions. Since the inner part of the propeller is stalled, it will create lots of drag for little thrust. Therefore, the max. RPM and torque numbers are very hard to predict, and measurements at that point will be of no value for predicting the behavior of the prop once it runs at its design advance ratio. Air density should be less of a problem - it will drive up dynamic pressure and, consequently, the power which is needed to run the propeller at a given speed.

Thank you so much! I kept seeing your name and answers when learning about propellers, and you've helped a lot in the past as well! Cheers

@ Peter Kampf I followed your link to your other answer and calculation therein. I encountered a problem, that I asked as a question here, and would appreciate your thoughts: https://aviation.stackexchange.com/questions/55633/propeller-efficiency-calculation-different-by-two-methods

@dumbpropnerd: Your example has 3 m/s airspeed and needs 11 m/s speed in the prop disk plane in order to work. With that massive acceleration, the usual simplifications of simple propeller theory break down (speed increase is small when compared to flight speed), so different equations will give different results. Test your prop at higher speeds and the differences will shrink.

Kampf Thanks! I suspected this had something to do with this being a simplified theory, as I had done my other analysis using induction factors taken into account, etc. I also realized that the simple propeller theory doesn't even take into account propeller characteristics, just the flight speed, thrust req'd, and air density, before making an efficiency calculation. How is this possible? I understand that its the ideal propeller efficiency, but how is this useful without taking into account propeller geometry?

@ Peter Kampf As an exercise, I used the Simple Prop Theory to calculate this ideal Froude efficiency under many scenarios. I found that for a given known system (C_drag, frontal area), the airspeed (and thus required thrust, from drag calculations) and air density had no impact in changing the ideal efficiency. This result is not intuitive to me. Does this make sense to you? Here are my calcs: https://ibb.co/kPPqzU

@dumbpropnerd: The propeller details are hidden in the prop efficiency. With a poor twist distribution (think fixed prop at the wrong speed, for example), this number will reflect all effects that are caused by the non-ideal nature of the prop. And what you describe sounds just like a fixed prop at the wrong speed.

@ Thank you. In that case, what utility do the calculations have? If you look at the image I posted, the results are using the Simple Prop Theory (no propeller chosen, just what the flight requires). Does this mean that for the given flight conditions, the propeller will be inefficient regardless of twist and prop characteristics?


text:
Propeller Efficiency calculation different by two methods
Please refer to this MIT page which I used for my calculations:
https://web.mit.edu/16.unified/www/FALL/thermodynamics/notes/node86.html
I calculated my aircraft propeller's prop, thrust, and torque curves, and subsequently the efficiency at a certain operating point (given the input thrust requirements and airspeed), by dividing the useful power (=Tu_0) by the input power (torqueRPS), where u_0 is the flight speed.
Now, in the propeller plane, the airspeed is u_disk according to Actuator Disk Theory, and can be calculated using the formulas and the known info. When I calculate the input propeller power according to another formula on that page [Power in = Thrust * u_disk], I get an answer not equal to the one using the torque*RPS formula. Where did this extra power go? I was expecting them to be the same, and the aerodynamic losses to be factored into the useful power being lower.

In my example, the flight speed is 3 m/s. Due to the conditions of the aircraft and environment, this requires 42N of thrust, which requires 4.5N-m of torque at 1503 RPM. This is all due to the propeller characteristics. Now, the useful work is T * airspeed = 42 x 3 = 126W.
Using the calculations on that page, we can get the accelerated speed at the disk is 11 m/s. So the input power (According to the page) is 42 x 11 = 464 W. This is different from the conventionally calculated input power of Torque * RPS = 4.5 * (1500/60) = 700 W.

Defining "useful" as moving the plane forwards, the rate at which useful work can be done is equal to whatever power which has been imparted to the fluid, i.e thrust x velocity at disc. This is propeller power. The rate at which useful work is done which is equal to whatever power which has gone into actually making the plane move forward, i.e thrust x freestream velocity (since freestream velocity and airspeed are interchangeable depending on the frame of reference). This is propulsive power
Consequently there are two different measures of efficiency, the propulsive efficiency which is a measure of how efficiently the plane uses the propeller power and the propeller efficiency which is a measure of how efficiently the propeller uses shaft power. The equation on the MIT webpage is for propulsive efficiency but you want to know how efficiently your shaft power (2π x torque x RPS) was used, therefore you have to calculate propeller efficiency so you don't use that equation.

Thanks. I meant "2*pi" which is why in my numerical calculation, the numbers work out. Sorry for the mistype, I couldn't get it to edit my comment.

However, in the MIT link I posted, it defines useful power out as thrust x upstream velocity u_0. That's my confusion. I think input power should be the same using the torque calculation and the velocity at disk calculation...

@dumbpropnerd sorry I misunderstood the definitons you were using, I amended my answer

Looked at your amended answer. That's actually the conclusion I initially arrived at, hence I asked "where does the extra power go" between the shaft, and the available power to be used on air (some of which goes to accelerate the air and some goes to generate thrust). But before it becomes "available power" why is it already lower than the shaft power? That's my confusion. I think that they should be the same...I am wondering if the discrepancy is from me using a 3D analysis to find the torque required, and the Disk Theory not accounting for 3D effects.

The extra power goes into creating tip vortices and swirling the flow, neither of which provide useful thrust therefore, that power is a loss.

I believe that is part of the loss, but the accelerated air is also part of the loss. According to this link I just found: https://s2.smu.edu/propulsion/Pages/efficiency.htm it shows the two methods of calculating efficiency I used, and equated them as being equal to each other. Thus, in my example, I should have 700 W for shaft power, 700W for thrust X disk velocity (assuming no shaft-losses) and then directly 126 W for useful work with no step down in the middle. From within that 700W, the losses towards accelerating air as well as swirl and vortices occur, I think.


text:
Propeller Characteristics at Different Altitudes
Assuming I have a propeller with a given diameter (D), pitch, blade twist, and number of blades, I can theoretically analyze discrete airfoil sections in XFOIL and come up with lift and drag polars for the airfoil sections, for a given Reynolds number. Using the new information along with the pitch and twist (to give angle of attack for each section), the overall thrust and torque for the propeller blade can be solved for given operating conditions (RPM and airspeed). Even more accurately, the induction airflow factors can be accounted for to adjust the velocity vectors and angles of attack experienced by the airfoil sections on the blade, and an iterative solution can yield more accurate thrust and torque data for the propeller. Doing this at different operating conditions (RPM and airspeed) can give many data points, allowing for a generation of a thrust, torque, and efficiency curve vs advance ratio. I have done all of the above for a certain propeller.
However, if I wanted to change my flight altitude (and thus air density would be different), my Reynolds number could be vastly different for the same operating conditions. Does the thrust, torque, and efficiency curves for the propeller still hold up? Or does the different Reynolds number mean that if I solved the airfoil sections thrust and torque data again, they would yield different dimensionless data as a function of advance ratio?
I'm happy to provide more details. Looking forward to figuring this out!

Using 2D airfoil data will give too conservative results. The slowing boundary layer is accelerated sideways by centrifugal forces, which means it is shifted to a section of the propeller blade that moves faster. As a result, flow separation is greatly delayed on propellers.

You need to look into blade element theory -- XRotor, JavaProp, and CCBlade are good examples.  Before the Reynolds number differences will be important, the density differences will dominate.  As you go higher, you will need to spin the (presumed fixed pitch) propeller at much higher RPM in order to make the same thrust.  Your power required to produce said thrust will also increase.  Look into the definitions of thrust and power coefficient and advance ratio.

Your lift and drag polar is Reynolds number dependent. You need to check if you are happy with what XFoil generates for you. Airfoil characteristics are fairly constant over a wide range of Re, so it might not effect you at all.
Also, if you are running high tip speeds, there is another pote increasing altitude drops the temperature as well, so you are increasing the tip Mach number, which you need to make sure stays within the validity limits of your calculation method.

Regardless of the Reynolds number, momentum theory says that if air density is decreased, the actuator disc has to move air faster in order to provide the same amount of thrust. Threfore the induced velocity across the actuator disc must increase. Going back to an actual propeller, a higher propeller RPM is required to provide this extra velocity therefore the required RPM will also rise.
Conversely, at a set RPM and induced velocity, less thrust will be produced. Since advance ratio is proportional to the ratio of free stream velocity to RPM, advance ratio would stay the same while the thrust at a given advance ratio would decrease. This would result in your advance ratio - thrust curves being "shifted down" along the Y axis.

So to clarify, propeller characteristic curves (e.g. thrust, torque, or efficiency vs Advance Ratio), are only valid at a certain density (/altitude)? At any change with density, they also change? Is there a way of quantifying this change without doing the XFOIL analysis from scratch, if the curves for one density are known?

@dumbpropnerd I use thin airfoil theory so that I don't have to manually punch in lift coefficient values from XFOIL for each blade element into my spreadsheet every time I have to re-sum the thrust and torque for the entire blade.

Thank you! I will look into this approach.

You said that the thrust-AR curve moves down. But wouldn't the torque-AR curve move down as well? In that case, is the propeller more or less efficient? It's producing less thrust, but also using less power to do so.

I guess so. This has nothing to do with efficiency, that's a tip vortex thing.

'tis a bit more interesting when it is fully considered. As density alters, to achieve a given thrust level, the AOA of the blade must alter, and while density is a common factor to both lift and drag, the CL/CD for varying AOA alters, a lot. That assumes the torque exists to turn the blender at all altitudes. The difference in AOA can be estimated, for a given thrust output the density delta compared to the base altitude will require a deltaCL to match. Easy. Except that the prop pitch is relative to TAS, so as TAS goes up, so does the geometric blade pitch, to achieve a constant AOA, so there is an additional delta for the tas, proportional to the tip velocity of the blade. As that increases the blade geometric angle to gain a given thrust, that results in a further increase in drag, and more annoyingly, a reduction in lift, as they are vectors dependent on... the relative air flow to the blade. Take it to the extreme, go fast enough, the blade needs to be near feathered angle, and it doesn't give much thrust then, but lordy it gives a force counter to the torque required to turn the whole blender thing.


text:
How does camber location (xf/c) affect Cl max and stall angle of an airfoil
I'm trying to compare 4-digit airfoil and 5-digit airfoil about the effect of camber location towards Cl max and stall angle.
Is it correct if i assume the higher the camber location value the Cl max and stall angle also increase. Also can anyone elaborate on why this is occur?

One can compare all kinds of designs at airfoiltools.com.
Higher camber tends to increase the "bending" of the airstream as angle of attack increases and is usually placed near the leading edge for this purpose, resulting in a higher coefficient of lift before stall.
Around the time when biplanes flew at 40 to 100 knots, thick wings with lots of camber proved to have a much wider flight envelope than the thinner wings of the time.  A comparison of the DAE21 cambered to the NACA 0012 symmetrical shows some of the advantages, including an almost 2x increase in Coefficient of Lift to Coefficient of Drag ratio.  In other words, the symmetrical needs to fly at 10 degrees AoA to produce the same lift as the cambered at 5 degrees.  Now the drag comparisons become interesting!
Improvements continued to be made up to the time of the Davis wing.  Only when increased airspeeds caused critical Mach number to become an issue did the age of these wings end, but retractable slats and Fowler flaps live on to help high subsonic airliners to slow down to 1/3 their cruising speed and land safely.
Cambered wings, such as the Clark Y, continue to be used with slower recreational aircraft.


text:
Flying wing airfoil
What would be let say some of appropriate airfoils for flying wing RC model ? Wingspan arround 1200mm and speed arround 700 km/h.
Maybe MH64 or so ?
Thank you,

I’m voting to close this question because it belongs on Drones.SE instead.

@Toni Bitenc 700 km/hr?

Tried migrating this 3 times, and it's just closed.  SE bug perhaps.

Why so many negative answers here? Could you explain please? I dont understand...:

700 km/h is very fast for rc aircraft, but not totally out of question. A couple of years ago, rc Jets exceeding 700 kmh/h (https://youtu.be/uCH64yciiq8) apeared

Just out of curiosity how is your project going to look like? :-)

Hello, for now it looks it wil be 60degree  triangle :-)) What to say :-)))

Challenging and therefore interesting project :-)  How are you going to control it? Horten airfoils are also definitely worth a look. Or, more generally, reflexed airfoils (but I suppose you already know this)

Thank you for your replay,   :-))) Headacke headacke :-))) For sure electronic will be needed to contro it for this speeds .

If you use reflexed airfoils it should remain self-stable. Does it have to fly so fast for a contest?

Hello, no contest just for fun

I recommend one that has been used successfully before, and at up to 1000 km/h airspeed: The Göttingen 765 (source).

Used as a root airfoil on the DfS 94 and 194 (aka Me 163 A). The root airfoil on the Me 163 B was 0.4% thinner, but otherwise very similar.

Thank you Peter
Will check also other topics as Im new here and will be back


text:
What airfoil family does turbofan's fan blades use?
The thing is, I am trying to make turbofan blades in Solidworks for a possible undergraduate thesis.
I want to know if there is any airfoil from the NACA series or any airfoil families plausible for this application. Maybe an airfoil in which I can reference the shape I will make because I would need to indicate the airfoil used in the thesis.
Any leads to design process involved in turbofan blades is appreciated, as this might help if no specific airfoils are available for the public. Thanks!

Hi Paul. I originally voted to close as needing more details. I see you added that this would be a University project. I still feel this is too vague. The theory on blade design is vast and detailed, to the point that there are entire books written on it. And even helping for the single specific project I think it would require an answer too long to be viable here. Are you just looking for a plausible airfoil? Or do you need to design a turbine for a specific engine/use?

I would like to know a plausible airfoil to be used for turbofan fan blade

Airfoil for the outer part of the blade is normally a simple convex profile (i.e. with pointy leading and trailing edge). Anyway the whole geometry is much more complex. I'd suggest you to get in touch either with some engine manufacturer or some writer of scientific article about turbofan simulation.

The blade shape is part of the proprietary art of jet engine design, which is a fiendishly complicated business. There are many engineering constraints and requirements which must be met for any one single engine design and it takes months and months of computer work to accomplish this.
But that's just the beginning of the blade design process, because after deciding on the blade shape, you have to look for a blade material that can handle all the different aerodynamic, mechanical, vibratory, thermal, and inertial stresses that will be applied to the blade while the engine is running. Optimizing the blade material requires years of materials science research and experimentation.
This is most assuredly not a task which can be undertaken by a single person in his or her garage!

Thanks, I am doing it for a university project in Solidworks though, not in real life. But I want to replicate or at least be able to make the general shape as accurate as possible from hub to tip. I'd be doing some CFD analysis on it as well, and it would be much appreciated if you can give me pointers on what to do, read or follow. Thanks!

I'm sorry, but I do not know any references to steer you to. Any good textbook on turbomachinery (as a branch of mechanical engineering) should furnish examples of the sort which may be helpful to you. good luck- Niels


text:
meanings of the numbers in the airfoil NACA 65(12)-10, where to find and plot
I am trying to plot an airfoil named NACA 65(12)-10 in javafoil. I found it in a few research articles. However, I am not sure from what series the airfoil is from. I tried looking it up on Airfoiltools website but to no avail. I am unsure if it is from the 6-digit series because it doesn't follow the naming nomenclature, such as NACA 65-210. Can anyone help me understand the name and how I can plot it? thanks.

There must be an error in the airfoil number. This one would have a width of the laminar bucket of 2.4 c$_L$, way too wide for any laminar airfoil. Also, for a proper 6-series name there need to be 3 digits past the hyphen, not two. Please explain where you found this odd number!

I found a research article link using the airfoil and a few more research articles relating to the topic link

According to this NACA report:
"In the designation NACA 65-(12)10, the first two digits, 65, refer to a particular thickness distribution; the second two digits, 12, indicate the isolated airfoil lift coefficient in tenths; and the third pair of digits, 10, denotes the maximum thickness in percent chord."
It seems that that airfoil is especially used for gas turbine compressor blades.

It looks like they're stretching the NACA definition -- this is fairly common, but it would be nice if they said exactly what they're doing in the paper or cited the paper that did.
Theory of Wing Sections by Abbott and von Doenhoff is your go-to reference for this kind of information.
Normally, a 65 series is numbered 65 XYY.  The X is the design lift coefficient in tenths and YY is the t/c in percent.  so, a 65310 would have a design cl=0.3 and t/c of 10%.
In this case, it looks like your airfoil has an extreme amount of camber such that the design cl=1.2 and it has a t/c of 10%.
I quickly generated something this in OpenVSP and it looks about right.  Right now, OpenVSP will only go up to cl=1.0.  I modified my local copy and generated this airfoil with cl=1.2 and t/c=10%.

I searched back through a few of the references that use this airfoil and this data set.  It appears that they've thickened the trailing edge and rounded the absolute end.  The theoretical airfoil in the above image has an extremely sharp cusped TE that would be impossible to manufacture.  Digging back further, the airfoil used becomes even less clear.


text:
Calculating True Course, Magnetic Heading and Compass Heading when Wind, True Heading, Magnetic Course and Deviation is known
I came across another exercise while studying for my PPL(A) exam where I would need some help with some calculations.
The question sounds like this:

Given: WCA: -012°; TH: 125°; MC: 139°; DEV: 002°E What are: TC, MH und
CH?
a.) TC: 113°  MH: 127°  CH: 129° b.) TC: 137°  MH: 127°  CH: 125°
<--- this is the correct answer c.) TC: 137°  MH: 139°  CH:
125° d.) TC: 113°  MH: 139°  CH: 129°

I think for calculating the True Course (TC) I need to subtract the Deviation Angle (DEV)
from the Magnetic Course (MC): TC = 139° - 2° = 137°. Why subtracting, well based on "West is Best, East is Least" principle... I hope :)
For the Magnetic Heading (MH) I need to subtract the Wind Correction Angle (WCA) from the Magnetic Course (MC): MH = 139° - 12° = 127°. Why subtracting, well because the exercise tells me it's from an -012° angle... I hope :)
But how do I get now the Compass Heading (CH)?
Are actually any of my assumption correct?
Later Edit:
OK, so i notice now after reading more that my assumption for TC is a false positive. I think, normally TH = TC + WCA, so in my case 125 = TC + -12. So TC must be 137 in order for TH to be 125°
Based on the same assumption, if my Magnetic Course (MC) is 139°, applying the WCA would mean that my Magnetic Heading (MH) must be 127°.
Now subtracting the Deviation angle from the MH, would mean the CH is 125°...

With this legend: [given], to find=?, <calculated>, *answer*:
Given: WCA: -012°; TH: 125°; MC: 139°; DEV: 002°E
Find: TC, MH, CH
          C  +  D --> M  -  V --> T
Course         [2]  [139]         ?
+ WCA   [-12]       [-12]       [-12]
Heading   ?    [2]    ?         [125]

Compass course given magnetic course:
          C  +  D --> M  -  V --> T
Course  <137>  [2]  [139]         ?
+ WCA   [-12]       [-12]       [-12]
Heading   ?    [2]    ?         [125]

Compass heading given compass course:
          C  +  D --> M  -  V --> T
Course   137   [2]  [139]         ?
+ WCA   [-12]       [-12]       [-12]
Heading <125>  [2]    ?         [125]

Magnetic heading required given wind
          C  +  D --> M  -  V --> T
Course   137   [2]  [139]         ?
+ WCA   [-12]       [-12]       [-12]
Heading *125*  [2]  <127>       [125]

True course given wind:
          C  +  D --> M  -  V --> T
Course   137   [2]  [139]       <137>
+ WCA   [-12]       [-12]       [-12]
Heading *125*  [2]  *127*       [125]

Answer:
          C  +  D --> M  -  V --> T
Course   137   [2]  [139]       *137*
+ WCA   [-12]       [-12]       [-12]
Heading *125*  [2]  *127*       [125]


text:
How can ice form in an OAT of +20°?
According to a question that I came across while studying for my EASA exam on Aircraft General Knowledge, "icing" is most likely to happen in an OAT between "Between -5°C and +20°C", but how can this be?
I've read in books and online that because of supercooled water droplets, the "freezing" point can go above 0°C, but I didn't came across any writing that goes as high as 20°C.
And here the question from the exam:

This sounds like carb icing, as opposed to airframe icing? Could you confirm which - carb icing is possible at relative high OAT's due to the venturi and fuel vapour dropping temperatures inside the carb body considerably. Its a very different concept to airframe icing

Please provide a reference that we can crosscheck to validate that EASA actually says this, because this seems like an error.

I updated my question with a screenshot from the exam. As you can see it's not mentioning if it refers to the carburator. But, indeed, the questions before this one were about the carburator.

As a courtesy, please wait at least 24 hours before choosing an answer so that all the people around the world can get a chance to read and answer your question.

Assuming that this does relate to carburettor icing, which I also believe it does, I was taught that carb icing is much more likely to occur with OATs around 20-25°C than it is with OATs around -5to0°C.  The explanation being that at lower temperatures, ice crystals form fully and are less 'sticky' so bounce off the metal surfaces.  Sort of how snow can be 'too cold' to form snowballs.  When you add in the factor of a closed throttle offering a greater surface area for the ice to hit, this is why you should always apply carb heating on descents, especially when coming in to land.

pilots, (1) it seems inconceivable the test wouldn't make clear "what sort" of icing was under discussion? or (2) is the notion of the test that the putative pilot should certainly understand that from context? or (3) is it just a crap test, in the sense that these days, say, high school math tests are usually incredibly badly written?

@Fattie - I suspect it is a combination of 2 and 3.  Given the OP said that the questions before refer to the carburettor, maybe the question writer 'forgot' to write this one in a stand-alone context.  However, expecting a trainee pilot to be very aware of the risks of carb icing is not unfair.  When coming in to land, the risks are increased as is the danger - because of the low altitude.  Trainee pilots fly a number of solo hours and if all they know is that they 'should' apply carb heat without understanding why, they are in danger of not appreciating the importance of it and forgetting.

EASA produced a pamphlet for GA Piston Engine icing. It states:

Carb icing is not restricted to cold weather. It will occur on warm
days if humidity is high, especially at low power settings. Flight
tests have produced serious icing at descent power when the air
temperature was above 25°C, even with relative humidity as low as 30%.
At cruise power, icing occurred at 20°C when relative humidity was 60%
or more. [...]

Nice find! I extended the quote a bit and included the chart, since it can be quite useful. Hope you don't mind.

Dang, that's a nice chart.  Someone does science and engineering graphics really well.

This answer gives a good authoritative source confirming that it can occur, but doesn’t explain how it can occur at such temperatures.

Water cannot exist in solid form at a temperature of 20°C and at pressures in the order of 1 atm. That question is most definitely talking about carburettor icing.
Within a carburettor, temperatures can be much lower than OAT. This is a result of:

Vaporisation cooling: As the fuel evaporates within a carburettor, it needs energy (heat) to do so. It takes this heat from the surrounding air, thereby cooling it down.

Low pressure: Within the carburettor venturi the pressure is reduced, and so is the temperature.

These two factors can easily produce a peak temperature drop of 20°C. So even though the OAT may be relatively high, icing can still occur within the carburettor.


text:
Can the ECS pipe/hose run through the aircraft fuel tank?
Hope you can help?
We are in the process of redefining aircraft fuel tank and there might be a possibility that we may need to run the ECS (Environmental control system) pipe through the fuel tank? Would that be not possible (due to its size in diameter, thermal insulation, etc)?
The same question goes for electrical wiring routes and the hydraulic lines? I understand that there are rules of segregating those, but what could happen if they are all installed in the tank?
Thanks!

Electrical circuitry running through fuel tanks has been known to cause at least one catastrophic disaster - TWA flight 800. A short circuit likely ignited fuel vapor in the central tank

The four-year NTSB investigation concluded with the approval of the Aircraft Accident Report on August 23, 2000, ending the most extensive, complex and costly air disaster investigation in U.S. history at that time.[7][8] The report's conclusion was that the probable cause of the accident was explosion of flammable fuel vapors in the center fuel tank. Although it could not be determined with certainty, the likely ignition source was a short circuit.[1]: xvi  Problems with the aircraft's wiring were found, including evidence of arcing in the fuel quantity indication system (FQIS) wiring that enters the tank.

So just to clarify, what I take away from this answer is, “Yes, you can run electrical circuitry through fuel tanks.”

@Jim It was specifically in answer to (emphasis mine) "I understand that there are rules of segregating those, but what could happen if they are all installed in the tank?"


text:
How would ATC try to establish contact with a VFR craft?
Example: I'm in a 172 tailed N1234A, on a pleasure hop between Hicks Field and Addison, circling around DFW International to the north, flying at 1500ft within DFW's Mode C veil under VFR (so I'm squawking 1200). The approach direction for the day also happens to be the north (this is common; prevailing winds in the area tend to be southerly). I'm staying outside DFW's airspace and therefore have not checked in with DFW Approach.
As I'm about to pass underneath "Area B" of DFW's space (which begins at 2000ft between 10 and 13nm from the airport sweeping across the northern approach), there is a 747 approaching low in the class B umbrella. It will pass above and in front of me, but still too close for their comfort, and so ATC wants to verify that I see the jumbo out my left window and perhaps suggest I descend.
Questions:

Was flying under the approach traffic of DFW my first mistake? Should I instead be heading around south (which would be a really big detour as I'd need to avoid Love Field's portion of Class B) or just contacting DFW Approach for clearance through Class B space?
Am I in the wrong for not having made prior contact with DFW Approach or Fort Worth Center in these circumstances, even though I'm operating VFR in class E/G space?
Should I even be this close to Class B in the first place (I'm 500 feet below it and 2 miles outside the surface footprint)? If not, what's a better altitude and range for me to have planned for? Outside 13NM from the airport, I'd be under Area H which starts at double my altitude, 3000 ft, but would also require me to enter or avoid Alliance's Class D space, and it'd take me quite a bit out of my way (since it's just for fun and logging hours, I might as well just touch and go at DTO if it'll be this much hassle to avoid DFW).
Would DFW or FW Center even bother trying to inform me of the situation, assuming I stay safely underneath Class B space and the 747 stays safely in it? One would think, being in contact with the 747, they could simply tell the jumbo to maintain minimum 2500 or 3000ft until 10nm out to maintain 1000ft minimum vertical separation, but my experience with boats says that if someone should yield right of way in this situation, it's the smaller craft (me).
Here are the title questions. How would DFW attempt to contact me? What frequency would I be expected to be tuned to? The emergency frequency is of course 121.5, but during normal flight along this path I could conceivably be listening to Fort Worth Center (122.65), DFW Approach (124.3), Addison ATIS or Approach, etc.
What would they say that I'd be expected to respond to, and how should I respond? I'm guessing they'd say something to the effect of "unknown VFR craft bearing 350 range 12 altitude 15, acknowledge and identify".

You mention several times that you are not "in" DFW's airspace in this scenario. That is incorrect. You may be outside of the charted Class B airspace but Regional Approach owns the surrounding Class G and Class E airspace as far away as LUD, WEA, 50F, F41, or 1F7—basically a square with each side 35-40 miles out from DFW proper.

Long story short, ATC has no reason to contact you but if they were to try they'd probably use 121.5 (see below). You're not in their airspace and the whole purpose of that airspace is to ensure separation, so as long as you remain clear of it they won't be concerned about you although they'll probably report your position to the 747.
As for your specific questions:

That's your decision as PIC and no one else can make it. But I wouldn't spend much time in a zone where I'm at high risk of a wake turbulence encounter.
No, there is no requirement for you to be in contact with ATC in that airspace so you're not wrong. It would be good to get flight following if possible, but that's optional and might not be available anyway.
It's hard to say what "too close" means and it could change depending on the time of day, if you're thinking about traffic. Again, this is something that you as PIC have to determine.
They won't try to contact you since you're outside the class B and they have no reliable way to contact you anyway. And considering wake turbulence again, "too close for their comfort" is probably more like too close for your comfort :-) Aircraft size isn't a consideration for right of way (see 14 CFR 91.113).
They won't for the reasons above. And they also can't know if you have a working radio, which frequency you're on etc. But they would probably try 121.5 because a current FDC NOTAM states that you should be monitoring it if you can:

ALL AIRCRAFT OPERATING IN UNITED STATES NATIONAL AIRSPACE, IF CAPABLE,
  SHALL MAINTAIN A LISTENING WATCH ON VHF GUARD 121.5 OR UHF 243.0.

The AIM also suggests and encourages all pilots to monitor it (e.g. section 6-2-4).

It's hard to say, the FAA's ATC instructions don't cover this situation (or I couldn't find it). But giving a specific bearing is unlikely (a bearing from what?) because you might not have the equipment to verify it so perhaps "12 miles north of DFW at 1500ft" would be more likely for the position part.

Of course, if ATC really, desperately, urgently needs to contact you - like if you fly straight through the class B without a clearance - then they can always request an interception. That would really ruin your day, so monitoring 121.5 if you can is probably a good idea.

This is a complete, informed, well-written answer. Check.

Re: "interception"; I take that to mean from an F-15 or F-16 out of JNAS Lake Worth with rocket-powered cranium-shredders pointed my way?

@KeithS Not necessarily, any police or other law enforcement helicopter that can match your speed will do. If they have a F16 lying around, that would work though. :)

FWIW, I’ve had ATC call me on their normal sector frequency in hopes I was monitoring. That seems to happen a lot now that they can see ADSB data.

Normally they wouldn't even try to get in contact with you.  It is easier and probably blocks less time on the frequency to talk to someone they have established contact with than to someone who they don't know which frequency he's on.  What you said in point 4 is correct, they don't have control over you as you're VFR and you're not even in their controlled airspace.  Bigger aircraft don't have right of way over smaller aircraft. FAR-AIM Sec 91.113
What they will probably say to the 747 is something like "American 123, traffic, 3 o'clock 2500 feet, same direction, type unknown."
If AA123 acknowledge that they see you, "American 123, maintain visual separation", otherwise "American 123, maintain 3000 feet for traffic"

Bigger aircrafts don't have right of way over a smaller aircraft.  Interesting.  Do you have a reference for this?  Common sense, and self preservation, suggest the opposite.

@Simon, added far aim reference

Thanks for the answer. So I just do what I do and the tower will ensure the 747 misses me. Just for clarification, the example message you give sounds like it's mixing alert information and instruction; I'd be heading east across the approach so I'd be coming across the 747's path, so not sure what "same direction" means here, and also 2500 feet would be the 747's minimum altitude to maintain vertical spacing if I'm at 1500. The class B space I'm flying under starts at 2000 so I'd be in some trouble if I were flying 2500.

@KeithS VFR/IFR vertical separation minima is 500 ft

In practice they'll ask at approach sector frequency, "Regional to N123AB, are you on frequency?" My side of DFW is south and the approach is Regional 135.975. Although it's very rare in DFW and more common in Austin where confused pilots are typically at 119.0 anyway.

Pondlife's answer is excellent, but to expand on #1, anytime you are flying under a class B/C shelf, it is a very good idea to be talking Approach. There are usually a lot of planes down there squeezed into a very small space, so an extra set of eyes is invaluable.
In this specific case, Addison's ATIS will say to contact Approach--not Tower--on initial contact due to the airspace shape and traffic levels. Since you're going to have to call them anyway, you might as well do it as soon as you're off Hicks. They will probably clear you through the Bravo and vector you over the top of DFW, possibly without you even asking, which also avoids your wake turbulence problem and saves you time/money. Everyone wins.


text:
How do pilots "turn an airport around" at nontowered fields?
Consider Homestead Airport, a fictional paved, non-towered airfield in the exurbs of Bigtown, USA that has become popular as a home field for recreational pilots. The airport's busy enough that addressing "Homestead traffic" on MULTICOM any given daylight hour isn't just talking to yourself; there will be at least one other plane in the pattern and a couple more on the ground, however there isn't enough traffic yet to justify hiring ATCOs to run the field. The sole runway is 13/31.
On the Saturday in question, there's a southeasterly wind at sunrise and so traffic starts the day using runway 13. However, the wind suddenly changes just after noon as a norther blows through, such that the five or six aircraft currently in the traffic pattern now need to use runway 31.
What is the standard procedure, if any, to establish the new traffic pattern at the airfield given the wind change? I'm especially curious as to how traffic already in a pattern for the now-downwind runway would be expected to maneuver to establish the new pattern, especially if the segmented circle or AFD indicate traffic should stay to one side of the airport regardless of runway direction (for instance a left-hand pattern on runway 13 but a right-hand on 31, to keep traffic away from a developed neighborhood to the southwest).
Bonus question: At the time a transition to the new runway direction is needed, there is an aircraft holding short for runway 13, with another behind him. These aircraft need to taxi down to the other end of the airfield, and quickly, or they'll be staring at the nosecones of exiting planes. There is a single taxiway, too narrow to 180 on especially with other aircraft using it, and the nearest runway exit is halfway down the field, so aircraft on the ground must taxi down the runway in the opposite direction to get to the correct end. What radio comms are recommended/required to advise traffic in the pattern that the runway has taxiing traffic?

This sounds a tad specific to have been made up on the spot... ;)

It's a question I've been thinking about for a while, since I first heard that pilots coordinate to do this. I'm most interested in whether there are any predefined maneuvers that pilots should know to manage the transition, like "all aircraft on downwind leg should make a one-minute 180 away from the airfield at the point where they'd normally turn base, then proceed downwind and turn base at the other end".

Your example is too specific to be useful to others, so let's consider the general case instead.
At an untowered field there is no "active" runway - they're all active.
There is usually a "preferred" runway - the one best aligned into the wind - but any runway can be used by any aircraft at any time, subject to whatever safety considerations the pilots mutually agree on. (At untowered fields with intersecting runways that have a lot of training on the field it's not uncommon to have traffic on the crosswind runway intersecting the preferred runway when students need to practice crosswind landings. It can be coordinated safely with a little cooperation between the pilots involved.)
The runway is "changed" by pilots announcing they're using a different runway.  There may also be a unicom operator at the field who announces a change in the preferred runway and lets incoming aircraft know about it.
Usually pilots who are already in the pattern for the "old" runway will complete their circuit and land, clearing the runway as expeditiously as possible.
Similarly pilots waiting at the threshold when the wind shifts may elect to either make their takeoff from where they are, or taxi around to the new preferred runway.  (In a case where the aircraft must enter the runway to turn around, or where they must back-taxi on the runway this obviously needs to be coordinated with the other pilots. The standard phraseology to use for this is "N12345 back-taxiing on runway XX" )

To your specific example, the directions of traffic patterns don't really matter much: As I mentioned aircraft already in the pattern will probably complete their circuit and land on the runway they were lined up for. Even in the case of Left Traffic for 13 and Right Traffic for 31 and two aircraft maneuvering to enter the pattern for opposite runways the standard pattern entry (45-to-downwind) will tend to keep airplanes apart (they won't ever be heading at each other in the downwind leg if they enter it on the 45, and it should at least ensure they see each other in the opposite-direction 45 and can coordinate to ensure they're never flying at each other). One airplane will probably have to extend their pattern to allow the other to clear the runway safely before they can land, however.

Traffic really should not really be arriving to both downwinds simultaneously anyway, particularly in your example with a single runway: Inbound aircraft should coordinate to arrive in the pattern for the preferred runway.

Aircraft already in the pattern have two choices: Land downwind if safe, or depart the pattern and re-enter it to land on the new preferred runway.  A downwind landing must obviously be coordinated with aircraft which may be preparing to depart in the opposite direction.
If landing downwind is unsafe (due to wind speed) or otherwise undesirable (e.g. due to other traffic) climbing above pattern altitude, leaving the area, and returning via a standard pattern entry for the opposite runway is usually the safest course of action.
Aircraft waiting to depart also have two choices: Depart downwind (which must be coordinated with aircraft arriving to the preferred runway), or taxi to the preferred runway end for takeoff. In cases such as you've described where back-taxiing is necessary this must be coordinated with pilots landing (in either direction) as the back-taxiing aircraft will be occupying the runway they need for landing.

The major complicating factor you didn't specifically mention, but which is implied by multiple aircraft "in the pattern" is closed traffic (aircraft making touch-and-go or full-stop/taxi-back landings and immediately taking off again).
Full-Stop/Taxi-Back landings are not a problem: Much like "regular" landing traffic they can land downwind, and upon exiting the runway taxi back to the preferred runway end for their next takeoff.
Touch-and-Go aircraft have two main choices: Land downwind as a full-stop and taxi around to the preferred runway end (which is probably the safest and most considerate thing to do), or complete a touch-and-go with the intention of departing the pattern and returning to set up for circuits to the new preferred runway.
Another important consideration are aircraft without radios (NORDO)- yes, these do exist). A visual scan of the traffic pattern is required at any uncontrolled field at any time to ensure you are not conflicting with NORDO aircraft.
These are often slower planes, so having one in the pattern landing "the wrong way" may require you to wait a while for them to complete their circuit before you can begin operations on the new preferred runway.

This is complex and unfortunately can create a dangerous situation. I fly out of KDYL and have seen a similar situation occur. Generally what I see is not that the wind changes but that there is no wind so both directions are used. The fact of the matter is, at an untowered airport the runway choice is up to you unless specific conditions are specified in the AF/D. The problem is that many people forget that or simply do whats easiest. At DYL the calm wind runway 05 which is great if you are departing to the east or north but I have seen plenty of people use 23 on calm days to depart to the south and west. Landings are generally justified similarly and people will enter the patter that best fits the flight path they are on. That being said when I fly in and I hear people calling legs of a particular pattern I will go for that as its the assumed pattern at the field. Remember it is your job to observe the flow of traffic as you would if you lost radio communications. If you feel the wind favors the pattern that is not being used you can wait until everyone is on the ground and announce your intentions for the other pattern.
In regards to your bonus question if you decide to taxi back (go down the runway to the other end) on the active you should not do so unless there are no approaching planes. Likewise once you have stated you have entered the active for a taxi back (and announced as such) no one else should be taking the active from the other end to take off. You should not be entering the active for taxi back if anyone is on approach unless the taxiway at the other end is free to leave the active in time. Thus you SHOULD NOT encounter other aircraft but people are people so that does not mean you WILL NOT encounter them.


text:
Did old propeller aircraft begin takeoff off-center to counter p-factor?
I have tried to ask this in chat and I have been invited to ask a full fledged question here.
Is it true that an old (say about between the two world wars) take-off technique, to compensate for the p-factor experienced at high throttle settings, was to put the aircraft misaligned to the runway and let the p-factor align the aircraft to the center-line?
Example: take an aircraft with a p-factor that makes it yaw to the left, the pilot would have placed the aircraft on the left side of the center-line, with the nose to the right (so, looking at the center-line).

Never heard of that method, but in his Taildragger Tactics book, Sparky Imerson describes how to land on a slight diagonal down the runway to help reduce crosswind component, so there's no telling how creative pilots can get when they have to.

An off-center takeoff gives the pilot a better view of the runway. Airplanes back then were almost all taildraggers, and forward visibility on the ground was poor.

Why ask in past tense? The Mü-23 motor glider has a central wheel and would need someone to hold a wingtip during take-off. Sometimes, a different technique is used: The aircraft is parked at a right angle to the (grass) runway, with the wingtip pointing into the intended take-off direction on the ground. When the pilot opens the throttle, he steps on the rudder, causing the aircraft to swerve into the desired direction and rotating into a horizontal attitude thanks to the centrifugal force.

Akaflieg München Mü-23 (picture source)

Why I ask in past tense? From the wiki page: Number built 1 not exactly common to come across, is it? :P

@Federico: Unless you are flying from the Königsdorf airfield.

Also, this aircraft seems to be doing it because of the single wheel, not to compensate for the p-factor. Are there examples of aircrafts/pilots doing it for p-factor related necessity?

@Federico: I admit, this is not answering your question. But p-factor is also at work when uprighting the Mü-23.

Peter, as you mention, this does not answer the question, but it is upvoted, meaning that the system behind SE thinks that my question is answered. Could you help me having a chance of receiving an answer?


text:
Hours needed to go from Private Pilot AMEL to ASEL
I was wondering about the minimum hours required to go from Private Multi-Engine Land to Private Single engine land.

Welcome to Aviation.SE! Since this might differ between jurisdictions, please add a corresponding tag (e.g. [easa-regulations] or [faa-regulations]).

For an FAA applicant.

3 hours of cross-country flight training in a single-engine airplane.
3 hours of night flight training in a single-engine airplane
3 hours of flight training in a single-engine airplane by reference to instruments
3 hours of flight training in a single-engine airplane in preparation for the practical test
10 hours of solo flight training in a single-engine

Some of these dual hours can be combined but this is straight from §61.109(a)

Is 3 hours of cross-country night flight training by reference to instruments in a single engine airplane sufficient?

@Someone Yes, that should work. Realistically, once you get 3 hours of the required instrument time, you will have more than 3 hours of night and cross country time since taxi, takeoff, and landings are not instrument time so you will end up with about 3.5 of those two if you did 3 hours exact of instrument under those conditions.


text:
Why cant all planes use the best engines like the GeNX or the Trent XWB?
Why do airline manufacturers not prefer using the best engines(max thrust: weight ratio and best efficiency) for their airplanes? Doesn't this just not lead to lower sales and losses for the company?
EDIT

I wish to know if all the planes currently in production could be replaced with the best engines available like the GeNX, Ge9X, or the Trent 1000. Would it not lead to better efficiency?

Boeing put a new engine on the 737 and called it the 737 MAX.  It didn't end well.

Why doesn't Airbus certify airplanes with the best engine in the market? For example, the A330 had a General Electric CF6-80E1 engine which was very small for it

Because when the A330 was first developed during the '70s, the General Electric CF-6 was the best engine then available. As easy as that. Newer versions of the A330 are equipped with more recent PW4000 or Trent 700. The latest iteration of the A330 is the "neo" version which is powered by RR Trent 7000, an engine which "received its EASA type certification on 20 July 2018".
EDIT TO THE EDIT IN THE QUESTION

I wish to know if all the planes currently in production could be replaced with the best engines available like the GeNX, Ge9X, or the Trent 1000

Updating something in the aerospace world takes time, money and a lot of paperwork. So much time, money and paperwork that deciding if it's better to update an old airframe or come up with a clean sheet design can be a difficult choice. Market-trend and competitors analyses help with this choice. And the A330 is actually a very good example of how difficult this choice can be: when the B787 was first revealed in 2004, Airbus decided that:

no, it was definitely no threat for the A330.
maybe yes, we can actually propose an improved A330 with some better aerodynamic and engines.
wait! What about updating the engines of the A380 instead?
yes ok, let's redesign the wing, the tailplane, the cockpit and use newer engines and call it A350.
yep, the B787 is definitely going to kill the A330 and the A350 is so far just a joke that  no airline is going to buy... let's do a clean sheet design and call it A350xwb.

Bingo! Or maybe not? The airlines actually liked and pushed for both the solution 2. and 5. and Airbus pursued both. The first one has given rise to the A330neo while the second one has resulted in the A350. Given the poor market results of the A330neo it might have been better to pursue the 3. instead of the 2. or maybe start sooner with just the 5., who knows...
Anyway, it is not always that difficult: the choice between continually updating (also with newer and newer engines) the bestselling A320 (last year 571 were produced, i.e. one every 15 hours) or replacing it with a new maybe-better clean sheet design, is a no-brainer.

Oh, I guess my question kinda diverted from what I wanted to ask...

@SambhavKhandelwal the math basically is quite simple. As Sophit described upgrading engines is a tedious process requiring planning, new parts for the plane, testing, certification etc. This all adds up to a huge amount of money, and the new engines are simply not efficient enough to make the swap profitable.

@Bianfable the A330neo has been so disappointing in terms of selling that I didn't even considered it as being its successor  I'll look for some more pertinent example

Note that the "neo" in Airbus' naming literally stands for "new engine option".

All commercial aircraft use the best engine technology available at the time when they were being designed. Once the design is settled (after years of work!) an engine change would reset the whole design process and delay the plane's introduction.
This means that airframe designers must work hand-in-glove with engine manufacturers (at times, more than one) to make certain that a new engine design will be ready for installation in a new airframe- to ensure the most up-to-date engine possible for the application. Note that the history of aviation is full of examples where the release to manufacturing of a new plane design was delayed by the unavailability of a promised engine.
There is a natural temptation to upgrade older airframes by re-engining them with the latest & greatest in engine technology, but this involves retrofitting a plane with 10,000 hours on it with brand new engines that won't keep the plane generating revenue long enough to pay off the investment.
So when re-engining occurs, it is mostly done on airframes that are still in production so a new engine goes into a new plane. Note also that the customer buying the plane can specify at the time of purchase which type of engine to install, that best suits their use model.

Quite a few airframes give the buyer a choice of engine, even for the initial production run. It can depend on the requirements of the customer.

Define "best". Best for what? If there were a single universal "best", nothing else would exist.
As already said, aircraft use engines that are available when the type gets designed (or which are designed in parallel with them). Reengining, while possible, is extremely involved and expensive and tends to not be worth the effort. In fact it may well cost more than replacing the aircraft.
And then there are many operators who will want all aircraft in their fleet to have as much parts commonality as possible. Which is why major manufacturers tend to offer their aircraft with a choice of engines. Say a major blue airline has a long history of operating aircraft with GE engines. They're more likely to buy your aircraft if you offer it with engines of that brand, even if engines made by RR would have slightly better performance (and define that, speed? fuel burn? hours between overhaul?) because in their infrastructure it is cheaper to operate those slightly less efficient engines.
Or you are operating an airline in a country that's under an embargo that prevents you from importing parts for a PW engine, but you can import parts for an RR engine. If your favourite Airbus is only offered with PW engines you now can't buy it.

In additional to the points raise in the other correct answers, 'best' cost wise does not mean most fuel efficient.
If you have a fleet that you have operated for a decade and know that after every x hours you need to service part y, part z can be done every other service and part q you just measure and replace as required then swapping to a 5% more fuel efficient engine that you need to change all of y,z and q four times as often while you learn how things wear may not gain anything. Labor and parts costs can easily overwhelm fuel savings.
Or you can work to the old processes and kill people. There is also a substantial safety risk from frequent change- The 737 Max crashes were in part driven by desire to update the 737 without changing training and processes. Regular engine updates will mean a long tail of incidents and near misses as previously highly skilled teams have to unlearn old information and learn the new, across maintenance, fueling/planning and actual flying, especially where efficiency has been gained at cost of complexity/performance.


text:
Why do commercial airliners go around on hard touchdown?
I was watching this video where he mentioned that a plane went around on hard touchdown. This isn't the only instance I have seen where they do this.
So, why do they do so? Isn't it a loss of profits for the airline?

Just s a note, it might be better to ask about a bounce after touchdown, a plane can make a (relatively) hard landing without bouncing, deliberately, if runway conditions require that. This technique is called "positive landing", and it will surprise the passengers if the pilot does not inform them of that before.

Thanks for the reply @Jpe61 but I was unable to get what you meant to say.

A hard touchdown does not usually lead to go around but a bounce (like the one shown in your video) does, or at least it shoul, as explained in the answers.

For what it's worth, violations of safety that lead to accidents are much larger losses to the airline. It is much cheaper to go around than pay for the consequences of even the least damaging accident (many listed in the answers below).

Tail Strike
In larger commercial airliners such as the Airbus A320 series, Boeing 737, or similar, a hard bounce during landing significantly increases the risk of a tail strike.
After a hard bounce, these aircraft tend to sink rapidly due to the low airspeed. This can create a sensation of the aircraft "falling," which may instinctively prompt the pilot to pull back on the controls. Alternatively, the pilot may just keep adding back pressure to the controls which they started during the flare, whereas the correct technique is to hold the landing attitude1. In both cases, doing so reduces tail clearance and increases the likelihood of a tail strike.
According to the Boeing 737 Flight Crew Training Manual (FCTM):

If the airplane bounces during a landing attempt, hold or re-establish
a normal landing attitude and add thrust as necessary to control the
rate of descent. Thrust need not be added for a shallow bounce or
skip. If a high, hard bounce occurs, initiate a go-around.

Tunisair A320 bounce with tail strike
Austrian A320 bounce with near tail strike
For the Boeing 737, this situation is further complicated because the initial bounce may have occurred due to the thrust being higher than idle, which can cause the speedbrakes to not actually deploy on touchdown. If the thrust levers are then retarded to idle during the bounce, automatic speedbrake deployment can occur. This leads to a loss of lift and a nose-up pitching moment, potentially resulting in a tail strike or a hard landing on the subsequent touchdown (Boeing FCTM).
“Porpoising”
For other aircraft types, such as high-wing turboprop aircraft, a bounce can cause these aircraft to be more prone to "flying again" after a bounce. The danger here lies in the pilot's potential reaction to push the nose down to force the aircraft onto the ground, which could lead to "porpoising" or a hard impact on the nose gear (or even the propeller itself on light aircraft). Please see @Cloud ‘s answer for more info.
Safety
While efficiency is important in aviation, safety is paramount. A go-around may result in time penalties and burn an additional 600 kg of fuel, but it ensures the safety of the flight. The flight may remain profitable despite a go-around and additional 15 mins of flying.

1: For a jet aircraft the controls still have good authority at landing speeds, whereas during training on lighter aircraft the controls feel very sloppy at such speeds. The pilots may instinctively go back to training muscle memory as opposed to the correct technique of holding the landing attitude. Alternately they may just keep pulling back as it is instinctive to reduce rate of descent, though this reduces tail clearance.

"The flight may remain profitable despite a go-around and additional 15 mins of flying": but since safety is paramount, whether the flight remains profitable or not isn't a factor in the decision.  Even if the flight ends up losing money, it will lose less money than would a crash.

Yes, if you are driving on the freeway/motorway and miss your exit, you aren’t going to start reversing your car backwards just to get to the exit. The safest thing to do continue to the next exit. Likewise in aviation a hard bounce is a potentially dangerous, unsalvageable situation.

Additionally the airplane can falsely assess its position as being on the ground when it  is  actually aloft.

Would it be fair to say that unless a runway is extremely long relative to a plane's requirements (e.g. double or more), a pilot who is suitably conscientious about deciding whether to continue a landing that doesn't go perfectly according to plan would generally use enough runway while making that decision that a goaround would be indicated for that reason even if for no other?

@supercat some situations require a mandatory go around - such as a hard bounce or windshear etc etc. Other situations such as long landing alerts may be ignored and the landing attempt continued if the runway is long enough. Depends on the situation you find yourself in. Also strongly depends on your company’s standard operating procedures (SOPs)

Here's a good reason. If you bounce hard enough, the plane flies again. To get it down, some people then push the nose down, which causes another hard landing and another bounce, this is know as "porpoiseing down the runway" and is a quick way to run out of runway and overshoot the threshold.
Profits "should" not be prioritized over safety, hence the go-around.

That video is wild!

For aircraft the size of an a320 series, a seen in the video, the plane will typically sink fast after a bounce rather, the greatest risk being a tail strike.

Boingy! Boingy! Boingy!

Isn't it a loss of profits for the airline?

It surely costs more than a perfect landing would have, but at the point of deciding to go around, a perfect landing is no longer an option.  As others have noted, the go-around is mandated on safety grounds, so the cost/benefit analysis requires looking at the consequences of not going around: the risk of damaging the aircraft, the cost of additional inspections to assess whether such damage has occurred, possible loss of the aircraft altogether, and liability to the airport, passengers, shippers of cargo, adjacent landowners, etc., in the event of damage or injury.  The cost of the go-around is justified on these grounds, as are the costs of safety measures generally.
In other words, not going around might result in a far greater loss of profit for the airline.